Current aircraft braking systems utilize either steel or carbon disks that serve as the friction materials and heat sinks. Steel-based systems were originally used as brake materials, until the emergence of carbon-carbon (C—C) composite materials in the 1970s. C—C composites are now the state-of-the-art material for aircraft brake heat sinks and are being used in the vast majority of new military and large commercial aircraft programs.
Ceramic Matrix Composites (CMCs) exhibit some extraordinary thermal and mechanical properties and hold the promise of being outstanding materials for aircraft brake friction applications, as well as attractive candidates for the next generation heat sink materials for such applications. A particular CMC system that indeed possesses the potential for use as a next generation aircraft brake material, and offers potential breakthrough performance, has recently been identified. In particular, brake materials based on a boron carbide matrix composite melt infiltration system have been shown to offer extremely attractive benefits relative to both steel and carbon brake materials.
As a class of materials, ceramics are known to possess low density in comparison to steel and other metal alloys, high hardness and high oxidation resistance; some of them also have attractive heat capacity and thermal conductivity. Compared to the C—C used today, ceramics have the potential of providing some key performance advantages in terms of reduced wear rate, enhanced oxidation resistance, and reduced heat sink weight and/or volume.
The earliest attempts to use using ceramics for aircraft braking system applications were based on monolithic ceramics and cermets. However, none of these attempts were successful. The major cause for the resultant failures was due to the inadequate mechanical properties, especially low impact resistance and low fracture toughness, in conjunction with the well-known characteristic brittleness of ceramics. Thus, activities on the next-generation heat sink materials for aircraft braking system applications have focused on the development of fiber-reinforced CMCs that would improve the fracture toughness and impact resistance (reducing the brittleness) while retaining the other advantages of ceramics.
The two prime candidate CMC material systems, identified for aircraft braking system heat-sink applications due to thermo-mechanical considerations, are based on silicon carbide (SiC) and boron carbide (B4C). Between these two material systems, the B4C-based CMCs have the particular attractions that B4C is the third hardest material known, with only diamond and cubic boron nitride being harder, and that it has a heat capacity greater than both SiC and carbon. However, B4C-based fiber-reinforced CMCs suitable for aircraft brake application had not been made due to processing difficulties associated with B4C. Generally, previous attempts to make B4C-based CMCs were limited to materials without fiber reinforcements. For example, U.S. Pat. No. 5,878,849 issued Mar. 9, 1999, describes a cermet material made by infiltrating a pressed preform of B4C powder (not filament or fiber) with aluminum.
Silicon-filled CMCs have been reported in both U.S. and foreign literature to show improved friction coefficients and/or wear life in certain configurations. See, for example: R. W. Froberg and B. A. Grider, “High Friction Carbon/Carbon Aircraft Brakes”, 40th Int. SAMPE Symp., May 8 11, 1995, extended abstracts, pp 942-944; R. W. Froberg and T. E. Pratt, “Brake System with Improved Brake Material”, U.S. Pat. No. 4,815,572 issued Jul. 24, 1987 (assigned to Parker Hannifin Corp); W. Krenkel, “CMC Materials for High Performance Brakes”, ISTA Conference on Supercars, Aachen, 31 Oct.-4 Nov. 1994 (paper from the German Aerospace Research Establishment Institute of Structures and Design), Stuttgart; A. Lacombe, “Friction System Using Refractory Composite Material”, U.S. Pat. No. 5,007,508 issued Apr. 15, 1991 (assigned to SEP, France).
In 1994, the German Aerospace Research Establishment reported sub-scale dynamometer results on a C—C+SiC composite which showed improved stability, lower wear, and shorter processing times than C—C materials. (See W. Krenkel, ibid.) Pyrolyzed resin impregnated carbon fiber preforms, infiltrated with Si at 1500° C., yielded composites containing ˜35% SiC by weight. Friction coefficients varied between 0.2 and 1.0, higher than for C—C under comparable conditions, increasing with decreasing velocity. Wear was not affected by temperature up to 900° C.
Lower net wear rates as compared to two carbon-carbon disks wearing against each other were disclosed in U.S. Pat. No. 5,007,508 covering aircraft brakes in which a C—C composite disk is worn against a disk containing carbon or SiC fibers and the CVI matrix consists of SiC as the principal phase with minor amounts of C or BN on the fibers.
While the foregoing examples illustrate the potential advantages of Si-based CMC's, few of these claims have been independently substantiated. Very often, the friction and wear (F&W) test duty cycle, including load, pressure, or length of testing time and number of cycles were either not reported, or were far less severe than those demanded under realistic aircraft braking conditions. Furthermore, many of these examples only cited either the friction or the wear results, by themselves, instead of the more relevant combined F&W data.
Certain CMC systems are disclosed in a series of U.S. patents issued to Singh et al., assigned to The General Electric Company. Those patents all disclosed the use, inter alia, of carbon fiber preforms, and in some embodiments the use of boron carbide is described. These General Electric patents fall generally into one of two groups. Some of these patents describe the production of solid state sintered ceramic bodies, wherein the composite matrix is densified by hot-press sintering, and where the final sintered body is reduced in size from the body before sintering. Those patents, generally unsuitable for producing complex, near-net shape ceramic matrix components, are: U.S. Pat. No. 4,886,682, “Process for Producing a Filament-Containing Composite in a Ceramic Matrix”, issued Dec. 12, 1989; U.S. Pat. No. 4,915,760, “Method of Producing a Coated Fiber-Containing Composite”, issued Apr. 10, 1990; U.S. Pat. No. 4,931,311, “Method of Obtaining a Filament-Containing Composite with a Boron Nitride Coated Matrix”, issued Jun. 5, 1990; U.S. Pat. No. 5,051,301, “Coated Fiber-Containing Composite”, issued Sep. 24, 1991; U.S. Pat. No. 5,067,998, “Fibrous Material-Containing Composite”, issued Nov. 11, 1991; U.S. Pat. No. 5,160,676, “Fibrous Material-Containing Composite”, issued Nov. 3, 1992; and U.S. Pat. No. 5,407,734, “Fiber-Containing Composite”, Apr. 18, 1995.
The other General Electric patents, as referred to above, disclose the infiltration of a porous body with a molten silicon infiltrant. These patents require that the fibrous materials are coated entirely with boron nitride to avoid a reaction and bonding between the silicon infiltrant and the fibrous material, and to retain fiber pull-out capabilities and fracture toughness. Those patents are U.S. Pat. No. 4,889,686, “Composite Containing Coated Fibrous Material”, issued Dec. 26, 1989; U.S. Pat. No. 4,944,904, “Method of Obtaining a Fiber-Containing Composite”, issued Jul. 31, 1990; U.S. Pat. No. 4,981,822, “Composite Containing Coated Fibrous Material”, issued Jan. 1, 1991; U.S. Pat. No. 5,021,367, “Fiber-Containing Composite”, issued Jun. 4, 1991; U.S. Pat. No. 5,043,303, “Filament-Containing Composite”, issued Aug. 27, 1991; U.S. Pat. No. 5,330,854, Filament-Containing Composite”, issued Jul. 19, 1994; U.S. Pat. No. 5,376,427, “Ceramic Composite Containing Coated Fibrous Material”, issued Dec. 27, 1994; U.S. Pat. No. 5,387,299, “Ceramic Composite Containing Coated Fibrous Material”, issued Feb. 7, 1995; and U.S. Pat. No. 5,432,253, “Composite Containing Fibrous Material”, issued Jul. 11, 1995.
In summary, prior to the present invention, there was no known practical processing technique for producing dense, fiber-reinforced B4C based CMCs having a fine grain microstructure and process for forming such composites. Preferably, the boron carbide particulates are maintained in sub-micron size in the CMC. Also, desirably, the growth of SiC crystals (grains) in the CMC above 10 microns is prevented, most preferably the growth of SiC crystals above 5 microns is prevented. The term micron as used herein is synonymous with the term micrometer.